Method for repairing an aircraft fuselage

ABSTRACT

A method for repairing damage to the coating of an aircraft fuselage, the coating being made of fiber-reinforced composite material. This method comprising making a cut-out in the fuselage with a substantially parallelepipedal outline around the damage and drilling a circular hole substantially centered on the intersection of the sides of the outline of the parallelepiped of the cut-out at each of the corners of the outline before the cutting out operation.

BACKGROUND OF THE INVENTION

The invention relates to a method for repairing aircraft primarystructures. More specifically, the invention relates to a method forrepairing damage to the coating of an aircraft fuselage, which coatingis made of fiber-reinforced composite material.

While in operation, aircraft structures are frequently subjected toimpacts with birds, vehicles, tools, airport installations, etc. Suchimpacts create damaged areas within the structure, which have lowermechanical resilience; these areas become prime sites for the start ofdefects such as cracks, which are likely to propagate.

More specifically, in the case of a fuselage made of fiber-reinforcedcomposite material, the actual damaged area may be significantly largerthan the visible impact area.

When an aircraft is subjected to such damage, it must absolutely berepaired so as to recover all the structural strength and avoid thepropagation of defects from the damaged area before it can be put backinto service.

It must be possible to carry out such repairs within the shortestpossible immobilization time for the plane.

A known repair mode involves covering the damaged area with a generallycircular liner or patch, whose area is significantly larger than saiddamaged area, then fastening it to the portion of the structure thatremains sound with any appropriate means, such as rivets, bolts, weldingor bonding.

This solution has the advantage of being relatively simple to implement.It does, however, have the disadvantage of keeping the damaged zone,which remains under loading and therefore may initiate the propagationof defects, even though the presence of the liner limits the flow offorces to which said area is subjected. Consequently, such repairs canonly be very temporary and need to be monitored very closely until apermanent repair can be effected.

In addition, the liner must be shaped to match the shape of the fuselagein the area under consideration. This shape may be complex andnon-involute, such that it requires specific shaping of the liner, whichbeing circular in shape must then be cut-out of a larger-size plate,itself made by any sheet metalworking means.

According to another embodiment described, for example, in internationalpatent application WO2007135318 in the name of the applicant, apolygonal cut-out is made around the impact area so as to eliminate thewhole of the damaged area. A liner, also polygonal in shape but with alarger surface area, is then fastened to the part that remains sound soas to close the cut-out. This method avoids the initiation of defects inthe damaged area, since this has been eliminated. Nevertheless, it isthen necessary to ensure that the cutting-out operation itself does notintroduce any defects. In the case of patent application WO2007135318,holes centered on the intersection of the sides of the polygon aredrilled at each corner of said parallelepiped before cutting-out tofacilitate this and to avoid unfortunate “saw-cuts” in the delicatecut-outs at the corners. The method divulged in this patent applicationaims therefore to avoid the formation of cutting defects which may giverise to cracks at the corners of the polygonal cut-out. Effectively, itis in these areas that the beginnings of cracks are most likely tooccur, subsequent to a lack of precision on the part of the operatorassigned to cutting-out.

In the case of a fiber-reinforced composite material fuselage, defectssuch as delamination can be caused during cutting-out, irrespective ofthe care taken by the operator. The action of the cutting implement, ofwhatever kind it may be, can easily break the cohesion of plies locatedat the edge of the cut-out and thus initiate delamination. Where thematerial that is cut-out is a fiber-reinforced composite material, therisk of introducing defects at the edge of the cut-out dependsessentially on the rate of advance of the tool. At a given cuttingspeed, too high a rate of advance favors the occurrence of delamination,whereas too slow a rate of advance causes thermal degradation of thematrix (burning or melting). This type of cutting-out operation infiber-reinforced composite materials must therefore be performed at acontrolled rate of advance, within a narrow band of admissible speeds.In the case of repair operations, however, the cutting-out advancemovement is generally communicated manually to the machine by theactions of the operator who moves it along the path to be cut. Even if ajudicious choice of tool geometry can increase the range of favorablecutting-out conditions, it remains difficult or even impossible toprevent delamination of the surface plies from occurring, except wheresophisticated methods of automatic advance are implemented, whichcontrol the cutting parameters and, in particular, the rate ofadvance/cutting speed combination along complex trajectories that matchthe shape of the fuselage.

There is therefore a requirement for a repair method for an aircraftfuselage made of composite material that can be implemented withadequate safety in conditions compatible with the means of airportmaintenance workshops and requiring the shortest possible immobilizationtime for the plane.

BRIEF SUMMARY OF THE INVENTION

To achieve this, the invention proposes a method for repairing a damagedarea of the fiber-reinforced composite coating of an aircraft fuselagecomprising the following steps:

-   -   realize a polygonal cut-out around the damaged area;    -   fit a coating liner with a polygonal outline, an area greater        than the cut-out and adjusted to the shape of the fuselage so as        to close the cut-out;    -   fasten said liner to the fuselage skin that remains sound, with        at least two rows of fasteners laid out parallel to the        perimeter of the liner;    -   a circular hole centered on the intersection of the sides of the        polygon having been cut-out in the fuselage at each of the        corners of the polygonal outline of the cut-out before the        cutting operation.

This method, implemented on a composite material fuselage, has twoadvantages:

-   -   prior cutting-out of the circular holes is essentially an axial        machining operation, which can be carried out by portable means,        giving precise control over the cutting conditions;    -   the sides of the polygon can be cut out manually, using any        appropriate means such as an angle grinder, jig-saw, portable        milling machine or router, therefore with less efficient control        over the cutting conditions, in particular the advance movement,        because, surprisingly, according to this embodiment, these areas        are not under loading in operation once the liner has been        fitted.

Thus, the cut-outs at the corners are free of defects because of thestrict control over the cutting conditions; as for cutting out thesides, even if they were to have defects, these would not be likely topropagate because of the weak stresses to which these areas aresubjected.

Advantageously, the diameter of the holes will be between 20 and 40 mm.This diameter is adequate, firstly, to remove the loading from the sidesof the cut-out over an adequate width while, secondly, still remainingcompatible with conventional drilling means, such as step-drills mountedon an automatic drilling machine (ADM) or portable orbital drillingsystems, that allow these holes to be cut out in a single operation,with a single tool and thus avoid the risks linked to adjustmentoperations.

According to an advantageous embodiment, the liner is not fastened tothe fuselage at the corners. Typically, when N rows of fasteners are tobe installed, those fasteners located on the outside of an n×n diagonalare not installed. Advantageously, the value of n is between 2 and N.

This embodiment allows for better distribution of the load over all thefasteners and therefore better dissipation of the forces into thestructure that remains sound.

According to an embodiment of the method which is the subject of theinvention, the liner is made of metal. Thus it will be easier to adjustto the shape of the fuselage using conventional metalworking means. Morespecifically, if the fuselage is made of carbon-fiber reinforcedcomposite material, the liner will advantageously be made of titaniumalloy, both for their galvanic compatibility and for theirclosely-matched coefficients of thermal expansion.

According to another embodiment, the liner is made of plies of fiberspre-impregnated with organic resin, manually laid-up, shaped directlyonto the fuselage on the area of application of the repair.

According to this embodiment, the shape of the liner adjusts itself tothe shape of the area it covers during the shaping operation.Advantageously, when the resin is a thermosetting type, curing can beentirely realized on the fuselage.

BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS

The invention will now be described more precisely in the context ofpreferred non-limiting embodiments shown in FIGS. 1 to 8 in which:

FIG. 1 represents a synopsis of the repair principle which is thesubject of the invention, with the drilling of the end holes in FIG. 1A,the cutting out in FIG. 1B and the fitting of the liner in FIG. 1C;

FIG. 2 shows a detail of FIG. 1 C, which highlights the distribution ofthe fasteners at the corners of the liner;

FIG. 3 illustrates the distribution of the fasteners at the corners ofthe liner in a general case;

FIG. 4 represents the distribution of the service stress in the fuselagein the vicinity of a repair according to the prior state of the art;

FIG. 5 represents the distribution of the service stress in the fuselagein the vicinity of a repair according to an example of realization ofthe invention;

FIG. 6 is an example of repair in a general case that requires cuttingout a longitudinal stiffener in front view;

FIG. 7 is a B-B cross-section view along the thickness of the repairshown in FIG. 6;

FIG. 8 is a transversal A-A cross-section view of this same view in theimmediate vicinity of the repair.

DETAILED DESCRIPTION OF THE INVENTION

All these figures represent an example of realization that implements aparallelepipedal cut-out and liner. The person skilled in the art willadapt these lessons to the more general case of a polygonal cut-out.

FIG. 1A: the coating of a composite fuselage (1) has been subjected toan impact (100). To perform the repair, the operator starts by drillingfour holes (10) using an axial machining device, which holes delimit arectangular area surrounding the impact. The dimensions of saidrectangular area are given by the plane's maintenance manual, dependingon the characteristics of the coating and on the characteristics of theimpact.

FIG. 1B: from these holes, the operator realizes the cut-out lines (11)to extract the recommended rectangular area from the coating. Thiscut-out is realized with a radial advance movement tool, to which theadvance movement is communicated manually by the operator. As anexample, for a fiber reinforced composite coating, this tool can consistof an angle grinder, a jig-saw or a router such as a portable millingmachine available under the brand name ONSRUD® manufactured by LMTTools.

A liner (20) is then applied so as to close the cut-out; it is fastenedto the portion of the fuselage that remains sound using N rows offasteners (30), 2 in this case.

FIG. 2: N=2 and so as to distribute the dissipation of forces into thesound portion, the fastener provided for at position (31) located on theoutside of the 2×2 diagonal (302) is not fitted at the corner of theliner. FIG. 3: generally, those fasteners located on the outer part ofthe corner of the liner are not fitted in one or several rows. In theexample of FIG. 3, N=4, the fasteners are not fitted at the locationsbeyond the 4×4 diagonal (304) i.e. 6 fasteners not fitted including the2×2 diagonal (302) and the 3×3 diagonal (303). The number of fastenersnot fitted at the corner depends on the number of rows and on theintensity of the loading, i.e. of the surface area of the damaged areaand its location on the fuselage. This configuration is determined bycalculation and consigned in the maintenance and repair manual for themost frequent cases.

FIGS. 4 and 5: comparing the services loading at the edge of the cut-outcan be performed using a finite element simulation. Both figurescorrespond to the same macroscopic loading conditions and give the VonMises yield criterion in the repaired fuselage in the vicinity of thecut-out. The reference loading, which corresponds to areas distant fromthe cut-out (110), is the same in both cases. The loading level for theother areas is defined relative to this reference loading.

FIG. 4: by using the cutting out principle of the prior state of the artwith skipped connections at the edges of the cut-out, the loading of thestraight parts of the cut-out is between 60% and 100% of the referenceloading.

FIG. 5: by using the cutting out principle according to the invention,the service loading at the straight edges of the cut-out is reduced by80%, compared with the reference loading in the immediate vicinity ofthis edge and still remains 40% lower when moving substantially awayfrom the edge; the isodynamic line (111, 112) delimits this 40%reduction in the reference loading which extend even beyond the firstrow of fasteners.

FIG. 6 represents the general case of a fuselage repair, in particularwhere the area to be cut-out goes through a longitudinal stiffener orstringer (40). In the case of a composite fuselage, such stiffeners arelinked to the skin by bonding, cocuring or other techniques of assemblywithout fasteners. In this case, the stringer is cut out over a width L,greater than the width of the cut-out. The width L is equal to the widthof the liner (20) and will therefore vary as a function of the number ofrows of fasteners required to fasten it to the sound part of thefuselage and to transmit the flow of forces.

FIG. 7: a shim (120) whose thickness is equal to the thickness of theskin is placed in the cut-out. An inner liner (121) whose thickness isdesigned to compensate the thickness of the stringer's flange (41) andwhose dimensions are substantially equal to those of the outer liner(20) is placed inside the fuselage. A splice plate (42) is fitted toensure the mechanical continuity of the stiffener (40).

Fasteners (30) are positioned along an appropriate number of rows, whichgo through and assemble the outer liner (20), the fuselage skin thatremains sound (1) and the inner liner (121). In the cut-out area, saidfasteners go through and assemble the outer liner, the shim (120) andthe inner liner.

Where the stringer (40) goes through in the vicinity of the cut-out,fasteners (32) go through and assemble the outer liner, the shim (120),the inner liner (121) and the flange of the stringer splice plate (42).

Outside the cut-out area, the same type of fastener (32) assembles andgoes through the outer liner, the fuselage skin, the inner liner and thebase of the stringer splice plate.

Lastly, FIG. 8: beyond the outer liner, a third type of fastener (33)assembles and goes through the outer skin, the flange (41) of theoriginal stringer and the flange of the stringer splice plate.

Advantageously, the stringer splice plate (41) is made of metal,preferably titanium alloy approximately 1 mm thick.

The fasteners can be of rivet type but are preferably of the boltedtype, such as Hi-Lite® fasteners, supplied by the company with the samename, which provide an assembly that is easily mounted and removed toperform a final repair.

The above description clearly illustrates that through its variousfeatures and their advantages the present invention realizes theobjectives it set itself. In particular, since the section of thecut-out is designed to significantly reduce the stresses at the edges ofthe cut-out, the composite aircraft fuselage repair method according tothe invention can be implemented with adequate safety in conditionscompatible with the means of maintenance workshops of airports andrequiring the shortest possible immobilization time for the plane.

The invention claimed is:
 1. A repair method for a damaged area in thecoating of a fiber-reinforced composite fuselage, comprising: drilling acircular hole, substantially centered on an intersection of sides of apolygonal outline of a cut-out in the fuselage around the damaged area,at corners of said outline; cutting lines of the cut-out that extendfrom the corners; fitting a coating liner with a polygonal outline, theliner having an area greater than an area of the cut-out and adjusting ashape of the fuselage so as to close the cut-out; fastening said linerto a fuselage skin outside of the damaged area with a plurality offasteners, each respective edge of a perimeter of the liner havingadjacent thereto N rows of plural fasteners, laid out parallel to therespective edge, wherein N is a positive integer greater than or equalto 2, and wherein at each respective corner of the liner, a diagonal ofthe fasteners is fitted, wherein a first end of the diagonal being aterminal fastener of an outermost row of plural fasteners on a firstedge of the coating liner and being N fasteners away from a second edgeof the coating liner, and a second end of the diagonal being is a secondterminal fastener of an outermost row of plural fasteners on the secondedge of the coating liner and being N fasteners away from the first edgeof the coating liner, and wherein fasteners are not fitted outside thediagonal in a direction toward the respective corner.
 2. The methodaccording to claim 1, wherein the diameter of the circular holepreviously cut out is between 20 and 40 mm.
 3. The method according toclaim 1, wherein each circular hole is cut-out in a single operation andwith a single tool.
 4. The method according to claim 1, wherein theliner is made of a metal alloy.
 5. The method according to claim 1,wherein the liner is made of a stack of pre-impregnated plies shapedwhen hot on the fuselage.
 6. An aircraft comprising a repair realizedaccording to claim
 1. 7. The method according to claim 1, wherein thecut-out is centered on a longitudinal stringer of the fuselage.
 8. Themethod according to claim 1, wherein the cut-out overlaps an elongatestiffener comprised in the fuselage, the stiffener having transverseflange portions connected by a middle stiffening portion, the methodcomprising: cutting the stiffener in a transverse direction so as toremove a longitudinal portion of the stiffener between two cut ends ofthe stiffener; providing in the cut-out a shim; providing inside thefuselage an inner liner that extends longitudinally between the cutflange portions of the stiffener; and fastening together the liner, theshim, and the inner liner inside of the cut-out area with a plurality ofthe fasteners.
 9. The method according to claim 8, wherein the shim hasa thickness equal to a thickness of the fuselage skin surrounding thecut-out.
 10. The method according to claim 8, wherein the inner linerhas a thickness substantially equal in thickness to the flange portionsof the stiffener.
 11. The method according to claim 8, wherein thefasteners inside the cut-out area are aligned with the N rows offasteners.
 12. The method according to claim 8, wherein the longitudinalportion is wider than the cut-out.
 13. The method according to claim 8,further comprising: providing a splice plate having transverse flangeportions connected by a middle stiffening portion, the splice platehaving a thickness and an elongate shape that extends between andoverlaps the two cut ends of the stiffener such that the flange portionsof the splice plate overlap the flange portions of the two cut ends ofthe stiffener to provide mechanical continuity across the fuselagebetween the two cut ends where the longitudinal portion of the stiffeneris removed, wherein the splice plate abuts the inner liner.
 14. Themethod according to claim 13, further comprising: fastening together theliner, the fuselage skin, the inner liner, and the splice plate outsideof the cut-out area with a plurality of second fasteners laid outlongitudinally in the flange portions of the splice plate.
 15. Themethod according to claim 13, further comprising: fastening together theliner, the shim, the inner liner, and the splice plate inside of thecut-out area with a plurality of second fasteners laid outlongitudinally in the flange portions of the splice plate.
 16. Themethod according to claim 13, further comprising: fastening together thefuselage skin, the stiffener, and the splice plate in areas where thesplice plate overlaps the stiffener with a plurality of third fastenerslaid out longitudinally in the flange portions of the splice plate. 17.The method according to claim 8, wherein the liner and the longitudinalportion of the stiffener removed from the fuselage are equal in width.18. The method according to claim 8, wherein the stiffener is linked tothe fuselage skin by bonding.
 19. The method according to claim 8,wherein the stiffener is linked to the fuselage skin by cocuring.